Abstract: A flow device for a combustor assembly defining a longitudinal axis around which the flow device is positioned is provided. The flow device includes a body extended around the longitudinal axis. The body includes a first wall and a second wall separated longitudinally from one another along the longitudinal axis. A fuel injector passage is defined through the body coaxial to the longitudinal axis. A plurality of vane walls is extended radially between the first wall and the second wall relative to the longitudinal axis. A radial flow passage is extended between the plurality of vane walls and the fuel injector passage.

Abstract: An apparatus and method for a combustor, the combustor including combustor liner having a plurality of dilution openings. The combustor receives a flow of fuel that is ignited and mixed with dilution air to form a flow of combustion gases. The flow of combustion gases travels through the combustor to a turbine section of an engine.

Abstract: The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

Abstract: A combustor assembly comprising a deflector wall in which a plurality of openings is defined through the deflector wall and around the fuel nozzle opening. The plurality of openings defines a first set of openings at a first radius, a second set of openings at or greater than a second radius greater than the first radius, and a third set of openings at one or more of a third radius between the first radius and the second radius. The first set of openings defines one or more of a first angle relative to the radial direction between approximately 60 degrees and approximately 100 degrees. The second set of openings defines one or more of a second angle between approximately zero and approximately 30 degrees. The third set of openings defines one or more of a third angle between the first angle and the second angle.

Abstract: The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

Abstract: A flow device for a combustor assembly defining a longitudinal axis around which the flow device is positioned is provided. The flow device includes a body extended around the longitudinal axis. The body includes a first wall and a second wall separated longitudinally from one another along the longitudinal axis. A fuel injector passage is defined through the body coaxial to the longitudinal axis. A plurality of vane walls is extended radially between the first wall and the second wall relative to the longitudinal axis. A radial flow passage is extended between the plurality of vane walls and the fuel injector passage.

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a liner extending between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and the aft end. The combustor assembly additionally includes a dome attached to or formed integrally with the liner, the dome and the liner together defining at least in part a combustion chamber. A fuel nozzle is attached to the dome, the fuel nozzle configured as a premix fuel nozzle for providing a substantially homogenous mixture of fuel and air to the combustion chamber, the mixture of fuel and air having an equivalence ratio of at least 1.5.

Abstract: A mixer assembly for a turbine engine is generally provided. The mixer assembly includes a vane assembly including a plurality of vanes configured to direct a flow of oxidizer to mix with a flow of fuel. The vane assembly includes a fluid diode disposed within a vane flow path between each pair of vanes of the vane assembly.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a deflector wall, an annular axial wall, and an annular shroud. The deflector wall is extended at least partially along a radial direction and a circumferential direction relative to an axial centerline and adjacent to a combustion chamber. A fuel nozzle opening is defined through the deflector wall, and a nozzle centerline is extended through the fuel nozzle opening along a lengthwise direction. The annular axial wall is coupled to the deflector wall and extended through the fuel nozzle opening. The axial wall is defined around the nozzle centerline. The annular shroud is defined around the nozzle centerline and extended co-directional to the axial wall. The axial wall and the annular shroud are each coupled to a radial wall defined upstream of the deflector wall. The annular shroud, the axial wall, and the radial wall together define a cooling plenum therebetween.

Abstract: A mixer assembly for a turbine engine is generally provided. The mixer assembly includes a vane assembly including a plurality of vanes configured to direct a flow of oxidizer to mix with a flow of fuel. The vane assembly includes a fluid diode disposed within a vane flow path between each pair of vanes of the vane assembly.

Abstract: A combustor assembly comprising a deflector wall in which a plurality of openings is defined through the deflector wall and around the fuel nozzle opening. The plurality of openings defines a first set of openings at a first radius, a second set of openings at or greater than a second radius greater than the first radius, and a third set of openings at one or more of a third radius between the first radius and the second radius. The first set of openings defines one or more of a first angle relative to the radial direction between approximately 60 degrees and approximately 100 degrees. The second set of openings defines one or more of a second angle between approximately zero and approximately 30 degrees. The third set of openings defines one or more of a third angle between the first angle and the second angle.

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. Additionally, the liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end and defines a forward section extending from the quench air jets to the dome. The dome and the forward section of the liner are configured to be cooled substantially by one or both of impingement cooling or convective cooling.

Abstract: An apparatus and method for a combustor, the combustor including combustor liner having a plurality of dilution openings. The combustor receives a flow of fuel that is ignited and mixed with dilution air to form a flow of combustion gases. The flow of combustion gases travels through the combustor to a turbine section of an engine.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a deflector wall, an annular axial wall, and an annular shroud. The deflector wall is extended at least partially along a radial direction and a circumferential direction relative to an axial centerline and adjacent to a combustion chamber. A fuel nozzle opening is defined through the deflector wall, and a nozzle centerline is extended through the fuel nozzle opening along a lengthwise direction. The annular axial wall is coupled to the deflector wall and extended through the fuel nozzle opening. The axial wall is defined around the nozzle centerline. The annular shroud is defined around the nozzle centerline and extended co-directional to the axial wall. The axial wall and the annular shroud are each coupled to a radial wall defined upstream of the deflector wall. The annular shroud, the axial wall, and the radial wall together define a cooling plenum therebetween.

Abstract: The present disclosure is directed to a method of operating a combustion system to attenuate combustion dynamics. The method includes flowing, via a compressor section, an overall supply of air to the combustion system; flowing, via a fuel supply system, an overall flow of fuel to the combustion system; flowing, to a first fuel nozzle of the combustion system, a first supply of fuel defining a richer burning fuel-air mixture at the first fuel nozzle; flowing, to a second fuel nozzle of the combustion system, a second supply of fuel defining a leaner burning fuel-air mixture at the second fuel nozzle; and igniting the richer burning fuel-air mixture and the leaner burning fuel-air mixture to produce an overall fuel-air ratio at a combustion chamber of the combustion system.

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. The liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end. The quench air jets include a plurality of primary stage air jets and a plurality of secondary stage air jets. The plurality of primary stage air jets are each spaced from the plurality of secondary stage air jets along the axial direction and together provide the combustion chamber with a quench airflow.

Abstract: A fuel-air mixer assembly that includes a mixer portion, and a flare cup portion coupled to the mixer portion. The flare cup portion includes a side wall that including an inlet opening and a discharge opening defined therein. The side wall is oriented such that the discharge opening is axi-asymmetrically shaped relative to a centerline of the fuel-air mixer assembly.

Abstract: A gas turbine engine combustor conical-flat heat shield includes an annular conical section extending upstream from and being integral with a flat section with a flat downstream facing surface which may be generally perpendicular to or canted with respect to a centerline. The flat section includes radially outer and inner edges at least one of which is circular and circumscribed about a centerline and circumferentially spaced apart clockwise and counter-clockwise radial edges having an origin on the centerline. A gas turbine engine combustor includes conical-flat heat shields in one or more circular rows arranged in a non-symmetrical or asymmetrical pattern. Two or more groups (A, B, C) of the conical-flat heat shields in the circular rows may be mounted on a domeplate and one or more of the conical-flat heat shields is different in one or more of the groups (A, B, C).

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. The liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end. The quench air jets include a plurality of primary stage air jets and a plurality of secondary stage air jets. The plurality of primary stage air jets are each spaced from the plurality of secondary stage air jets along the axial direction and together provide the combustion chamber with a quench airflow.

Abstract: A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. Additionally, the liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end and defines a forward section extending from the quench air jets to the dome. The dome and the forward section of the liner are configured to be cooled substantially by one or both of impingement cooling or convective cooling.